Centrifugal compressor diffuser passage boundary layer control

ABSTRACT

A centrifugal compressor diffuser ( 42 ) includes a plurality of diffuser flow passages ( 22 ) extending through an annular diffuser housing ( 20 ) and circumferentially bounded by diffuser vanes ( 23 ) and axially bounded by forward and aft walls ( 101,   100 ). A diffuser boundary layer bleed ( 96 ) for the passages may include boundary layer bleed apertures ( 106 ) or slots ( 130 ) disposed through the forward wall ( 101 ) and a downstream facing wall ( 142 ) canted at an acute cant angle to a downstream diffuser airflow direction ( 103 ) in the passages. Diffuser bleed flow ( 112 ) is bled from a diffuser boundary layer. Boundary layer bleed apertures can be located downstream of throat sections ( 28 ) of the flow passages near pressure sides of the vanes. A centrifugal compressor ( 18 ) may include the diffuser surrounding an annular centrifugal compressor impeller ( 32 ) and apparatus for flowing impeller bleed flow ( 102 ) from a radial clearance between an impeller tip ( 36 ) and a diffuser annular inlet ( 27 ) with diffuser bleed flow either mixed or separately to cool a turbine ( 16 ).

GOVERNMENT INTERESTS

This invention was made with government support under governmentcontract No. W911-W6-11-2-0009 by the Department of Defense. Thegovernment has certain rights to this invention.

BACKGROUND OF THE INVENTION Technical Field

The present invention relates to bleed air from gas turbine enginecentrifugal compressors.

One type of gas turbine engine includes a centrifugal compressor havinga rotatable impeller to accelerate and, thereby, increase the kineticenergy of air flowing therethrough. A diffuser is generally locatedimmediately downstream of and surrounding the impeller. The diffuseroperates to decrease the velocity of the air flow leaving the impellerand transform the energy thereof to an increase in static pressure,thus, pressurizing the air.

A conventional gas turbine engine typically includes a compressor,combustor, and turbine, both rotating turbine components such as blades,disks and retainers, and stationary turbine components, such as vanes,shrouds, and frames routinely require cooling due to heating thereof byhot combustion gases. Cooling of the turbine, especially the rotatingcomponents, is important to the proper function and safe operation ofthe engine. It is known to bleed cooling air from the centrifugalcompressor to help cool the turbine.

Failure to adequately cool a turbine disk and its blading, for example,by providing cooling air deficient in supply pressure, volumetric flowrate or temperature margin, may be detrimental to the life andmechanical integrity of the turbine. Depending on the nature and extentof the cooling deficiency, the impact on engine operation may range fromrelatively benign blade tip distress, resulting in a reduction in enginepower and useable blade life, to a rupture of a turbine disk, resultingin an unscheduled engine shutdown.

Balanced with the need to adequately cool the turbine is the desire forhigher levels of engine operating efficiency which translate into lowerfuel consumption and lower operating costs. Since turbine cooling air istypically drawn from one or more stages of the compressor and channelledby various means, such as pipes, ducts, and internal passageways to thedesired components, such air is not available to be mixed with fuel,ignited in the combustor and undergo work extraction in the primary gasflowpath of the turbine.

Total cooling flow bled from the compressor is a loss in the engineoperating cycle and it is desirable to keep such losses to a minimum.

Some conventional engines employ clean air bleed systems to cool turbinecomponents in gas turbines using an axi-centrifugal compressor as isdone in the General Electric CFE738 engine. The turbine cooling supplyair exits the centrifugal diffuser through a small gap between thediffuser exit and deswirler inner shroud. Other turbine cooling airmethods include extracting cooling from the impeller or from a gapbetween the impeller and the diffuser exit.

U.S. Pat. No. 5,555,7211 to Bourneuf, et al, which issued on Sep. 17,1996 and is entitled AGas Turbine Engine Cooling Supply Circuit@,discloses using bleed air from an impeller stage of a centrifugalcompressor in a turbine cooling supply circuit for a gas turbine. U.S.Pat. No. 5,555,721 discloses impeller tip forward bleed flow andimpeller tip aft bleed flow for cooling turbine components. U.S. Pat.No. 5,555,721 is assigned to the General Electric Co., the same assigneeas this patent and is incorporated herein by reference.

U.S. Pat. No. 8,087,249 to Ottaviano, et al. which issued Jan. 3, 2012,and is entitled ATurbine Cooling Air From A Centrifugal Compressor@discloses a gas turbine engine turbine cooling system including animpeller and a diffuser directly downstream of the impeller and a bleedfor bleeding clean cooling air from downstream of the diffuser. U.S.Pat. No. 8,087,249 is assigned to the General Electric Co., the sameassignee as this patent and is incorporated herein by reference.

Thus, there continues to be a demand for advancements in diffuser designand geometry that improves aerodynamic performance and reduces theoverall engine radial envelope.

BRIEF DESCRIPTION OF THE INVENTION

A diffuser for a centrifugal compressor includes an annular diffuserhousing, diffuser vanes axially extending between a forward wall and anaft wall of the diffuser housing, a plurality of diffuser flow passagesextending through the housing and spaced about a circumference of thehousing. The diffuser flow passages are bounded by the diffuser vanesand the forward and aft walls. A diffuser boundary layer bleed isprovided for bleeding diffuser bleed flow from a diffuser boundary layerin each of the diffuser flow passages.

The diffuser boundary layer bleed may be configured for bleeding thediffuser bleed flow from the diffuser boundary layer at a positionlocated in a region of flow weakness in each of the diffuser flowpassages.

The diffuser boundary layer bleed may include boundary layer bleedapertures disposed through the forward wall. Each of the boundary layerbleed apertures may be a slot including a downstream facing wall angledor canted at an acute cant angle with respect to a downstream diffuserairflow direction in each of the diffuser flow passages respectively.

The boundary layer bleed apertures may be positioned or locateddownstream of throat sections of the diffuser flow passages nearpressure sides of the diffuser vanes.

A centrifugal compressor including an annular centrifugal compressorimpeller, a diffuser annularly surrounding the impeller, and a pluralityof diffuser flow passages extending through a housing of the diffuserand spaced about a circumference of the housing. Each of the passagesincludes a throat section and a diffusing section downstream of thethroat section. The diffuser flow passages are circumferentially boundedby diffuser vanes extending axially between forward and aft walls of thediffuser and a diffuser boundary layer bleed is provided for bleedingdiffuser bleed flow from a diffuser boundary layer in each of thediffuser flow passages.

The centrifugal compressor may also include a radial clearance betweenan impeller tip of the impeller and an annular inlet of the diffuser, ameans for mixing impeller bleed flow from the radial clearance withdiffuser bleed flow from the boundary layer bleed apertures forproviding turbine cooling air, and a means for flowing the turbinecooling air to a turbine or a means for flowing impeller bleed flow andthe diffuser bleed flow separately to the turbine.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 is a sectional view illustration of a gas turbine engine with acentrifugal compressor for mixing impeller tip bleed flow and diffuserbleed flow in the compressor section before using the flows for coolingturbine components.

FIG. 2 is an enlarged sectional view illustration of the centrifugalcompressor and a diffuser with diffuser bleed holes illustrated in FIG.1.

FIG. 3 is an aft looking forward perspective view illustration of thediffuser and the diffuser bleed holes through 3-3 in FIG. 2.

FIG. 4 is an enlarged perspective view illustration of the bleed holesillustrated in FIG. 3.

FIG. 5 is a perspective view illustration of a portion of the diffuserand the diffuser bleed holes illustrated in FIG. 2.

FIG. 6 is an enlarged sectional view illustration of the centrifugalcompressor tip and the diffuser bleed holes illustrated in FIG. 2.

FIG. 7 is a sectional view illustration of a gas turbine enginecentrifugal compressor with an alternative arrangement for separatelyflowing impeller tip bleed for cooling turbine components.

FIG. 8 is a sectional view illustration of the gas turbine engineillustrated in FIG. 7 with an arrangement for separately flowingdiffuser bleed flow for cooling turbine components.

FIG. 9 is an enlarged perspective view illustration of one of theimpeller bleed flow ports illustrated in FIG. 7 and as taken through 9-9in FIG. 10.

FIG. 10 is a forward looking aft perspective view illustration of an aftcasing surrounding the centrifugal compressor and including the impellerand bleed flow ports illustrated in FIGS. 7 and 8 respectively.

FIG. 11 is cutaway perspective view illustration of impeller bleedflowpaths for one of the impeller bleed flow ports illustrated in FIGS.7 and 9.

FIG. 12 is an enlarged perspective view illustration of one of thediffuser bleed flow ports illustrated in FIG. 8 and as taken through12-12 in FIG. 10.

FIG. 13 is cutaway perspective view illustration of a diffuser bleedflowpath through one of the diffuser bleed flow ports illustrated inFIG. 8 and as taken through 12-12 in FIG. 10.

DETAILED DESCRIPTION OF THE INVENTION

Illustrated in FIG. 1 is a gas turbine engine high pressure centrifugalcompressor 18 in a high pressure gas generator 10 of a gas turbineengine 8. The high pressure centrifugal compressor 18 is a finalcompressor stage of a high pressure compressor 4. The high pressure gasgenerator 10 has a high pressure rotor 12 including, in downstreamserial or flow relationship, the high pressure compressor 14, acombustor 52, and a high pressure turbine 16. The rotor 12 is rotatablysupported about an engine axis 25 by bearings in engine frames notillustrated herein.

The exemplary embodiment of the high pressure compressor 14 illustratedherein includes a five stage axial compressor 30 followed by thecentrifugal compressor 18 having an annular centrifugal compressorimpeller 32. Outlet guide vanes 40 are disposed between the five stageaxial compressor 30 and the single stage centrifugal compressor 18.Compressor discharge pressure (CDP) air 76 exits the impeller 32 andpasses through a diffuser 42 annularly surrounding the impeller 32 andthen through a deswirl cascade 44 into a combustion chamber 45 withinthe combustor 52. The combustion chamber 45 is surrounded by annularradially outer and inner combustor casings 46, 47. Air 76 isconventionally mixed with fuel provided by a plurality of fuel nozzles48 and ignited and combusted in an annular combustion zone 50 bounded byannular radially outer and inner combustion liners 72, 73.

The combustion produces hot combustion gases 54 which flow through thehigh pressure turbine 16 causing rotation of the high pressure rotor 12and continue downstream for further work extraction in a low pressureturbine 78 and final exhaust as is conventionally known. In theexemplary embodiment depicted herein, the high pressure turbine 16includes, in downstream serial flow relationship, first and second highpressure turbine stages 55, 56 having first and second stage disks 60,62. A high pressure shaft 64 of the high pressure rotor 12 connects thehigh pressure turbine 16 in rotational driving engagement to theimpeller 32. A first stage nozzle 66 is directly upstream of the firsthigh pressure turbine stage 55 and a second stage nozzle 68 is directlyupstream of the second high pressure turbine stage.

Referring to FIG. 1, the compressor discharge pressure (CDP) air 76 isdischarged from the impeller 32 of the centrifugal compressor 18, usedto combust fuel in the combustor 52, and to cool components of turbine16 subjected to the hot combustion gases 54; such as, the first stagenozzle 66, first and second stage shrouds 71, 69 surrounding the firstand second high pressure turbine stages 55, 56 respectively. The highpressure compressor 14 includes a compressor aft casing 110 and adiffuser forward casing 114 as more fully illustrated in FIGS. 1 and 2.The compressor aft casing 110 generally surrounds the axial compressor30 and the diffuser forward casing 114 generally surrounds thecentrifugal compressor 18 and supports the diffuser 42 directlydownstream of the centrifugal compressor 18. The compressor dischargepressure (CDP) air 76 is discharged from the impeller 32 of thecentrifugal compressor 18 directly into the diffuser 42.

Referring to FIGS. 2 and 3, the impeller 32 includes a plurality ofcentrifugal compressor blades 84 radially extending from a rotor discportion 82. Opposite and axially forward of the compressor blades 84 isan annular blade tip shroud 90. The shroud 90 is adjacent to blade tips86 of the compressor blades 84 defining a blade tip clearance 80therebetween. The diffuser 42 includes an annular diffuser housing 20having a plurality of tangentially disposed diffuser flow passages 22extending radially therethrough, spaced about a circumference 26 of thehousing 20, and through which diffuser airflow 103 flows in a downstreamdirection. Diffuser vanes 23 axially extend between a forward wall 101and the aft wall 100 of the diffuser 42.

Referring to FIGS. 2 and 3, the diffuser vanes 23 circumferentiallyextend between adjacent ones of the diffuser flow passages 22. Thediffuser flow passages 22 are partly defined and circumferentiallybounded by the circumferentially spaced apart diffuser vanes 23.Adjacent ones of the passages 22 intersect with each other at radiallyinner inlet sections 24 of the passages 22 that define a quasi-vanelessannular inlet 27 of the diffuser 42. Each passage 22 further includes athroat section 28 downstream of and integral with the inner inletsection 24. Each passage 22 further includes a diffusing section 99immediately downstream of the throat section 28.

Referring to FIGS. 2 and 6, a centrifugal compressor first cooling airsource 92 for turbine cooling air 88 is a small predetermined radialclearance (C) located between an impeller tip 36 of the rotatingimpeller 32 and the annular inlet 27 of the static diffuser 42. Impellerbleed flow 102 from the radial clearance (C) is collected in a radiallyinner manifold 104. The predetermined. radial clearance (C) is designedto accommodate thermal and mechanical growth of the impeller 32 and isopen to or in fluid communication with the radially inner manifold 104.

Referring to FIGS. 3-6, we have found that the diffuser airflow 103 onone side of the passage (such as passage 22) in multi-passage diffusers(such as the diffuser 42) that follow or are downstream of centrifugalimpellers (such as the impeller 32) is often weak and may be subject toseparation. Separation in the passage can generate high losses thatlowers engine specific fuel consumption (SFC). This area or region ofweak flow 127 is also believed to be a contributor to surge that limitsflow range of the compressor.

A centrifugal compressor stage second cooling air source 94 for turbinecooling air 88 includes a diffuser boundary layer bleed 96 for bleedingdiffuser bleed flow 112 from a diffuser boundary layer 113 in each ofthe diffuser flow passages 22 of the diffuser 42, illustrated herein asplurality of boundary layer bleed apertures 106. The diffuser boundarylayer bleed 96, also referred to as fluidic bleed, helps reduce the weakflow and limit or prevent the unwanted flow separation. The diffuserboundary layer bleed 96 bleeds diffuser bleed flow 112 from the diffuserboundary layer 113 into a radially outer manifold 116.

The radially inner and outer manifolds 104, 116 are in fluidcommunication such that the impeller bleed flow 102 from the radiallyinner manifold 104 flows into the radially outer manifold 116. Theimpeller and diffuser bleed flows 102, 112 are mixed in the radiallyouter manifold 116 to provide the turbine cooling air 88 which is thenported or otherwise flowed from radially outer manifold 116 through aplurality of circumferentially distributed manifold ports 117 to thehigh pressure turbine 16. The turbine cooling air 88 may be channelledor flowed therefrom by external piping (not shown) to cool the first andsecond stage shrouds 71, 69 (illustrated in FIG. 1).

Substantially axially extending beams or struts 122 separate theradially inner and outer manifolds 104, 116 and the impeller bleed flow102 passes between the struts 122 as it flows from the radially innermanifold 104 into the radially outer manifold 116. The fluidic bleedflow illustrated herein as the diffuser boundary layer bleed 96represents a small amount of flow, less than 1% of the engine core flow.The fluidic bleed is strategically removed near the inception of theweak flow to improve the overall performance of the diffuser.

Referring to FIGS. 3-5, the boundary layer bleed apertures 106 may beholes or slots 130 through the forward wall 101 of the diffuser 42 asillustrated herein. The boundary layer bleed apertures 106 or slots 130lead into and are in flow communication with the radially outer manifold116. The slot 130 is positioned or located downstream of the throatsection 28 near a pressure side 126 of the diffuser vane 23 at aposition where the flow would begin to show weakness or instability in adiffuser without the diffuser boundary layer bleed 96. This position islocated in what is referred to as a region of flow weakness 127. A slotwidth W may be sized with manufacturing constraints such as a minimumtool size. A slot length L may be selected to enable up to 3% of theengine core flow to be used.

The slot 130 should ideally be angled such that the diffuser bleed flow112 exits the slot perpendicular to a forward surface 105 of the forwardwall 101 of the diffuser 42 in a radial plane 132 passing through theengine centerline or axis 25 as illustrated in FIG. 5. However, becauseof constraints such as the slot extending through or very near a bend134 in the forward wall 101 of the diffuser 42 this angle may bedifferent. The slot 130 has radially outer and inner walls 136, 138, asillustrated in FIG. 6, and upstream and downstream facing walls 140,142, as illustrated in FIGS. 4 and 5 respectively, extending through theforward wall 101. The downstream facing wall 142 is designed to scoopboundary layer air 144 in the diffuser boundary layer 113 only. Thus,the downstream facing wall 142 is angled or canted at an acute cantangle B of less than 90 degrees with respect to the diffuser airflow 103(parallel to the direction boundary layer air 144 in the downstreamdirection in the diffuser flow passages 22 of the diffuser 42. Itappears that an acute cant angle B of 45 degrees is desirable. However,the acute cant angle B is limited by geometry and manufacturingconstraints on the outside of the diffuser so that an acute cant angle,for example about 22.5 degrees, is more practical.

Illustrated in FIGS. 7-13 is a gas turbine engine with a centrifugalcompressor similar to the one illustrated in FIGS. 1-3 but with analternative arrangement or design for separately gathering and flowingthe impeller tip bleed and diffuser bleed flow for cooling turbinecomponents. The impeller bleed flow 102 front the radial clearance (C),illustrated in FIG. 9, is flowed into and collected in a radially innerannular manifold 154 illustrated in FIGS. 7 and 9. Inter-manifoldapertures 160 are disposed between the inner annular manifold 154 and aplurality of radially outer annular manifolds 156 illustrated in FIGS.7, 9, and 13. The inter-manifold apertures 160 allow the impeller bleedflow 102 to flow front the inner annular manifold 154 into the outerannular manifolds 156. The impeller bleed flow 102 from the outerannular manifolds 156 is then ported or otherwise flowed through aplurality of circumferentially distributed impeller bleed flow manifoldports 157, illustrated in FIG. 10, to the high pressure turbine 16 forturbine cooling.

Referring to FIGS. 8, 10, and 11-13, the diffuser boundary layer bleed96 bleeds diffuser bleed flow 112 from the diffuser boundary layer 113into an annular diffuser bleed manifold 158 from where the diffuserbleed flow 112 is then ported or otherwise flowed through a plurality ofcircumferentially distributed diffuser bleed manifold ports 159 to thehigh pressure turbine 16 for turbine cooling. FIG. 10 illustrates therelative circumferential and axial locations of the impeller bleed flowmanifold ports 157 and the diffuser bleed manifold ports 159 on andthrough the diffuser forward casing 114.

While there have been described herein what are considered to bepreferred and exemplary embodiments of the present invention, othermodifications of the invention shall be apparent to those skilled in theart from the teachings herein and, it is therefore, desired to besecured in the appended claims all such modifications as fall within thetrue spirit and scope of the invention. Accordingly, what is desired tobe secured by Letters Patent of the United States is the invention asdefined and differentiated in the following claims.

What is claimed:
 1. A gas turbine engine centrifugal compressor diffusercomprising: an annular diffuser housing, diffuser vanes axiallyextending between a forward wall and an aft wall of he diffuser housing,a plurality of diffuser flow passages extending through the housing andspaced about a circumference of the housing, the diffuser flow passagesbounded by the diffuser vanes and the forward and aft walls, and adiffuser boundary layer bleed for bleeding diffuser bleed flow from adiffuser boundary layer in each of the diffuser flow passages.
 2. Thediffuser according to claim 1 further comprising the diffuser boundarylayer bleed configured for bleeding the diffuser bleed flow from thediffuser boundary layer at a position located in a region of flowweakness in each of the diffuser flow passages.
 3. The diffuseraccording to claim 1 further comprising the diffuser boundary layerbleed including boundary layer bleed apertures disposed through theforward wall.
 4. The diffuser according to claim 3 further comprisingeach of the boundary layer bleed apertures being a slot including adownstream facing wall angled or canted at an acute cant angle withrespect to a downstream diffuser airflow direction in each of thediffuser flow passages respectively.
 5. The diffuser according to claim3 further comprising the boundary layer bleed apertures positioned orlocated downstream of throat sections of the diffuser flow passages nearpressure sides of the diffuser vanes.
 6. The diffuser according to claim5 further comprising each of the boundary layer bleed apertures being aslot including a downstream facing wall angled or canted at an acutecant angle with respect to a downstream diffuser airflow direction ineach of the diffuser flow passages respectively.
 7. A gas turbine enginecentrifugal compressor comprising: an annular centrifugal compressorimpeller, a diffuser annularly surrounding the impeller, a plurality ofdiffuser flow passages extending through a housing of the diffuser andspaced about a circumference of the housing, each of the passagesincluding a throat section and a diffusing section downstream of thethroat section, the diffuser flow passages circumferentially bounded bydiffuser vanes extending axially between forward and aft walls of thediffuser, and a diffuser boundary layer bleed for bleeding diffuserbleed flow from a diffuser boundary layer in each of the diffuser flowpassages.
 8. The centrifugal compressor according to claim 7 furthercomprising the diffuser boundary layer bleed configured for bleeding thediffuser bleed flow from the diffuser boundary layer at a positionlocated in a region of flow weakness in each of the diffuser flowpassages.
 9. The diffuser according to claim 7 further comprising thediffuser boundary layer bleed including boundary layer bleed aperturesdisposed through the forward wall.
 10. The centrifugal compressoraccording to claim 9 further comprising each of the boundary layer bleedapertures being a slot including a downstream facing wall angled orcanted at an acute cant angle with respect to a downstream diffuserairflow direction in each of the diffuser flow passages respectively.11. The centrifugal compressor according to claim 10 further comprisingthe boundary layer bleed apertures positioned or located downstream ofthroat sections of the diffuser flow passages near pressure sides of thediffuser vanes.
 12. The centrifugal compressor according to claim 11further comprising each of the boundary layer bleed apertures being aslot including a downstream facing wall angled or canted at an acutecant angle with respect to a downstream diffuser airflow direction ineach of the diffuser flow passages respectively.
 13. The centrifugalcompressor according to claim 9 further comprising: a radial clearancebetween an impeller tip of the impeller and an annular inlet of thediffuser, a means for mixing impeller bleed flow from the radialclearance with the diffuser bleed flow from the boundary layer bleedapertures for providing turbine cooling air and flowing the turbinecooling air to a turbine, or a means for flowing the impeller bleed flowand the diffuser bleed flow separately to the turbine.
 14. Thecentrifugal compressor according to claim 13 further comprising each ofthe boundary layer bleed apertures being a slot including a downstreamfacing wall angled or canted at an acute cant angle with respect to adownstream diffuser airflow direction in each of the diffuser flowpassages respectively.
 15. The centrifugal compressor according to claim13 further comprising the boundary layer bleed apertures positioned orlocated downstream of throat sections of the diffuser flow passages nearpressure sides of the diffuser vanes.
 16. The centrifugal compressoraccording to claim 15 further comprising each of the boundary layerbleed apertures being a slot including a downstream facing wall angledor canted at an acute cant angle with respect to a downstream diffuserairflow direction in each of the diffuser flow passages respectively.17. The centrifugal compressor according to claim 9 further comprising:a radial clearance between an impeller tip of the impeller and anannular inlet of the diffuser, the radial clearance in fluidcommunication with a radially inner manifold, the boundary layer bleedapertures in flow communication with a radially outer manifold, theradially inner manifold in fluid communication with the radially outermanifold such that the impeller bleed flow flows into the radially outermanifold and mixes with the diffuser bleed flow to form turbine coolingair, and means for flowing turbine cooling air out of the radially outermanifold.
 18. The centrifugal compressor according to claim 17 furthercomprising each of the boundary layer bleed apertures being a slotincluding a downstream facing wall angled or canted at an acute cantangle with respect to a downstream diffuser airflow direction in each ofthe diffuser flow passages respectively.
 19. The centrifugal compressoraccording to claim 18 further comprising the boundary layer bleedapertures positioned or located downstream of throat sections of thediffuser flow passages near pressure sides of the diffuser vanes. 20.The centrifugal compressor according to claim 19 further comprising eachof the boundary layer bleed apertures being a slat including adownstream facing wall angled or canted at an acute cant angle withrespect to a downstream diffuser airflow direction in each of thediffuser flow passages respectively.
 21. The centrifugal compressoraccording to claim 9 further comprising: a radial clearance between animpeller tip of the impeller and an annular inlet of the diffuser, theradial clearance in fluid communication with a radially inner annularmanifold, inter-manifold apertures disposed between the inner annularmanifold and a plurality of radially outer annular manifolds, a meansfor porting and flowing the impeller bleed flow from the radialclearance through a plurality of circumferentially distributed impellerbleed flow manifold ports in and through an diffuser forward casingsurrounding the centrifugal compressor to the high pressure turbine forturbine cooling, the diffuser boundary layer bleed in fluid flowcommunication with and operable for bleeding the diffuser bleed flowinto an annular diffuser bleed manifold, and a means for porting andflowing the diffuser bleed flow through a plurality of circumferentiallydistributed diffuser bleed manifold ports in and through the diffuserforward casing to the high pressure turbine for turbine cooling.
 22. Thecentrifugal compressor according to claim 21 further comprising each ofthe boundary layer bleed apertures being a slot including a downstreamfacing wall angled or canted at an acute cant angle with respect to adownstream diffuser airflow direction in each of the diffuser flowpassages respectively.
 23. The centrifugal compressor according to claim22 further comprising the boundary layer bleed apertures positioned orlocated downstream of throat sections of the diffuser flow passages nearpressure sides of the diffuser vanes.
 24. The centrifugal compressoraccording to claim 23 further comprising each of the boundary layerbleed apertures being a slot including a downstream facing wall angledor canted at an acute cant angle with respect to a downstream diffuserairflow direction in each of the diffuser flow passages respectively.